EP4063062A1 - Procédés de réparation et systèmes de structures en nid d'abeille dans des moteurs à turbine à gaz - Google Patents

Procédés de réparation et systèmes de structures en nid d'abeille dans des moteurs à turbine à gaz Download PDF

Info

Publication number
EP4063062A1
EP4063062A1 EP22155874.5A EP22155874A EP4063062A1 EP 4063062 A1 EP4063062 A1 EP 4063062A1 EP 22155874 A EP22155874 A EP 22155874A EP 4063062 A1 EP4063062 A1 EP 4063062A1
Authority
EP
European Patent Office
Prior art keywords
braze
honeycomb structure
component
cover
pulling
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Pending
Application number
EP22155874.5A
Other languages
German (de)
English (en)
Inventor
Guolin Oo
Garimella Balaji Rao
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
RTX Corp
Original Assignee
Raytheon Technologies Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Raytheon Technologies Corp filed Critical Raytheon Technologies Corp
Publication of EP4063062A1 publication Critical patent/EP4063062A1/fr
Pending legal-status Critical Current

Links

Images

Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
    • B23P—METAL-WORKING NOT OTHERWISE PROVIDED FOR; COMBINED OPERATIONS; UNIVERSAL MACHINE TOOLS
    • B23P6/00—Restoring or reconditioning objects
    • B23P6/002—Repairing turbine components, e.g. moving or stationary blades, rotors
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
    • B23P—METAL-WORKING NOT OTHERWISE PROVIDED FOR; COMBINED OPERATIONS; UNIVERSAL MACHINE TOOLS
    • B23P6/00—Restoring or reconditioning objects
    • B23P6/002—Repairing turbine components, e.g. moving or stationary blades, rotors
    • B23P6/005—Repairing turbine components, e.g. moving or stationary blades, rotors using only replacement pieces of a particular form
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
    • B23K—SOLDERING OR UNSOLDERING; WELDING; CLADDING OR PLATING BY SOLDERING OR WELDING; CUTTING BY APPLYING HEAT LOCALLY, e.g. FLAME CUTTING; WORKING BY LASER BEAM
    • B23K1/00—Soldering, e.g. brazing, or unsoldering
    • B23K1/0008—Soldering, e.g. brazing, or unsoldering specially adapted for particular articles or work
    • B23K1/0014—Brazing of honeycomb sandwich structures
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
    • B23K—SOLDERING OR UNSOLDERING; WELDING; CLADDING OR PLATING BY SOLDERING OR WELDING; CUTTING BY APPLYING HEAT LOCALLY, e.g. FLAME CUTTING; WORKING BY LASER BEAM
    • B23K1/00—Soldering, e.g. brazing, or unsoldering
    • B23K1/0008—Soldering, e.g. brazing, or unsoldering specially adapted for particular articles or work
    • B23K1/0018—Brazing of turbine parts
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
    • B23K—SOLDERING OR UNSOLDERING; WELDING; CLADDING OR PLATING BY SOLDERING OR WELDING; CUTTING BY APPLYING HEAT LOCALLY, e.g. FLAME CUTTING; WORKING BY LASER BEAM
    • B23K1/00—Soldering, e.g. brazing, or unsoldering
    • B23K1/008—Soldering within a furnace
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
    • B23K—SOLDERING OR UNSOLDERING; WELDING; CLADDING OR PLATING BY SOLDERING OR WELDING; CUTTING BY APPLYING HEAT LOCALLY, e.g. FLAME CUTTING; WORKING BY LASER BEAM
    • B23K1/00—Soldering, e.g. brazing, or unsoldering
    • B23K1/20—Preliminary treatment of work or areas to be soldered, e.g. in respect of a galvanic coating
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
    • B23K—SOLDERING OR UNSOLDERING; WELDING; CLADDING OR PLATING BY SOLDERING OR WELDING; CUTTING BY APPLYING HEAT LOCALLY, e.g. FLAME CUTTING; WORKING BY LASER BEAM
    • B23K3/00—Tools, devices, or special appurtenances for soldering, e.g. brazing, or unsoldering, not specially adapted for particular methods
    • B23K3/08—Auxiliary devices therefor
    • B23K3/087—Soldering or brazing jigs, fixtures or clamping means
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
    • B23K—SOLDERING OR UNSOLDERING; WELDING; CLADDING OR PLATING BY SOLDERING OR WELDING; CUTTING BY APPLYING HEAT LOCALLY, e.g. FLAME CUTTING; WORKING BY LASER BEAM
    • B23K31/00—Processes relevant to this subclass, specially adapted for particular articles or purposes, but not covered by only one of the preceding main groups
    • B23K31/02—Processes relevant to this subclass, specially adapted for particular articles or purposes, but not covered by only one of the preceding main groups relating to soldering or welding
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
    • B23K—SOLDERING OR UNSOLDERING; WELDING; CLADDING OR PLATING BY SOLDERING OR WELDING; CUTTING BY APPLYING HEAT LOCALLY, e.g. FLAME CUTTING; WORKING BY LASER BEAM
    • B23K35/00—Rods, electrodes, materials, or media, for use in soldering, welding, or cutting
    • B23K35/02—Rods, electrodes, materials, or media, for use in soldering, welding, or cutting characterised by mechanical features, e.g. shape
    • B23K35/0222—Rods, electrodes, materials, or media, for use in soldering, welding, or cutting characterised by mechanical features, e.g. shape for use in soldering, brazing
    • B23K35/0244—Powders, particles or spheres; Preforms made therefrom
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/28—Supporting or mounting arrangements, e.g. for turbine casing
    • F01D25/285—Temporary support structures, e.g. for testing, assembling, installing, repairing; Assembly methods using such structures
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/005—Repairing methods or devices
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16J—PISTONS; CYLINDERS; SEALINGS
    • F16J15/00—Sealings
    • F16J15/44—Free-space packings
    • F16J15/444—Free-space packings with facing materials having honeycomb-like structure
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
    • B23K—SOLDERING OR UNSOLDERING; WELDING; CLADDING OR PLATING BY SOLDERING OR WELDING; CUTTING BY APPLYING HEAT LOCALLY, e.g. FLAME CUTTING; WORKING BY LASER BEAM
    • B23K2101/00—Articles made by soldering, welding or cutting
    • B23K2101/001—Turbines
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
    • B23K—SOLDERING OR UNSOLDERING; WELDING; CLADDING OR PLATING BY SOLDERING OR WELDING; CUTTING BY APPLYING HEAT LOCALLY, e.g. FLAME CUTTING; WORKING BY LASER BEAM
    • B23K2101/00—Articles made by soldering, welding or cutting
    • B23K2101/02—Honeycomb structures
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
    • B23K—SOLDERING OR UNSOLDERING; WELDING; CLADDING OR PLATING BY SOLDERING OR WELDING; CUTTING BY APPLYING HEAT LOCALLY, e.g. FLAME CUTTING; WORKING BY LASER BEAM
    • B23K2103/00—Materials to be soldered, welded or cut
    • B23K2103/18—Dissimilar materials
    • B23K2103/26—Alloys of Nickel and Cobalt and Chromium
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/001—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between stator blade and rotor
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/12—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • F01D11/127—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part with a deformable or crushable structure, e.g. honeycomb
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/23—Manufacture essentially without removing material by permanently joining parts together
    • F05D2230/232—Manufacture essentially without removing material by permanently joining parts together by welding
    • F05D2230/237—Brazing
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/80—Repairing, retrofitting or upgrading methods
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/55—Seals

Definitions

  • the present disclosure relates to repair systems for gas turbine engines, and more specifically, to methods and systems for adhering braze tape to a honeycomb structure.
  • a gas turbine engine typically includes a fan section, a compressor section, a combustor section, and a turbine section.
  • a fan section may drive air along a bypass flow path while a compressor section may drive air along a core flow path.
  • air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases.
  • the hot combustion gases flow through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads.
  • the compressor section typically includes low pressure and high pressure compressors, and the turbine section includes low pressure and high pressure turbines.
  • the turbine section includes multiple stages of blades and vanes. As fluid flows through the turbine section, the flow causes the blades to rotate about an axis of rotation.
  • the vanes, positioned between each row of blades, are used to redirect the flow in order to maximize the power received by the downstream blades.
  • honeycomb structures are used to provide gas-path seal between moving parts.
  • Honeycomb structures may also be used as sacrificial materials which rotating blades rub against the honeycomb structure while maintaining a predetermined tip clearance during engine operation.
  • a method of applying a braze component to a honeycomb structure is disclosed herein.
  • the method may comprise: applying at least a partial vacuum within a chamber, the chamber defined at least partially by a vacuum device and a cover, the honeycomb structure disposed within the chamber, the braze component disposed between the honeycomb structure and the cover; pulling the cover towards the braze component in response to applying the partial vacuum; and pulling the braze component into a plurality of hexagonal cells defined by the honeycomb structure in response to pulling the cover towards the braze component.
  • the braze component may comprise a braze tape.
  • the method may further comprise disposing the honeycomb structure on a porous platform prior to applying the partial vacuum.
  • the porous platform may be disposed within the vacuum device prior to applying the partial vacuum.
  • the method may further comprise disposing the braze component on the honeycomb structure, the honeycomb structure disposed between the porous platform and the braze component.
  • the method may further comprise creating an air-tight seal in the chamber between the cover and the vacuum device.
  • the honeycomb structure may be a portion of a honeycomb seal land for a sealing system in a turbine section of a gas turbine engine.
  • the method may further comprise heating the cover to soften the braze component while pulling the cover.
  • a method of repairing a honeycomb seal land of a turbine vane assembly may comprise: pulling a braze component into a plurality of cells of a honeycomb structure in response to applying at least a partial vacuum within a chamber, the braze component and the honeycomb structure disposed on a porous platform within the chamber; tack welding a second end of the honeycomb structure to an internal surface of an inner diameter platform of a vane assembly, the second end opposite a first end, the second end proximate the braze component; and brazing the braze component to couple the honeycomb structure to the internal surface.
  • the braze component may be a braze tape.
  • a vacuum device and a cover may at least partially define the chamber.
  • the honeycomb structure may be disposed between the porous platform and the cover prior to pulling the braze component.
  • the braze component may be disposed between the cover and the honeycomb structure prior to pulling the braze component.
  • the method may further comprise heating the cover during pulling the braze component.
  • the method may further comprise a weight on the cover prior to pulling the braze component.
  • a system for applying a braze tape to a honeycomb structure for use in a gas turbine engine may comprise: a vacuum device; a cover configured to create an air-tight seal with the vacuum device and define a chamber therein; and a porous platform disposed within the chamber, the porous platform configured to receive the honeycomb structure thereon.
  • the cover is configured to be heated when the system is in operation.
  • the system may further comprise a controller in operable communication with the vacuum device, the controller configured to command the vacuum device to supply at least a partial vacuum.
  • the porous platform may extend from a first side of the vacuum device to a second side of the vacuum device.
  • the system may be configured to pull the braze tape into a plurality of cells of the honeycomb structure.
  • any reference to attached, fixed, connected, or the like may include permanent, removable, temporary, partial, full, and/or any other possible attachment option.
  • Any reference related to fluidic coupling to serve as a conduit for cooling airflow and the like may include permanent, removable, temporary, partial, full, and/or any other possible attachment option.
  • any reference to without contact (or similar phrases) may also include reduced contact or minimal contact.
  • Cross hatching lines may be used throughout the figures to denote different parts but not necessarily to denote the same or different materials.
  • aft refers to the direction associated with the exhaust (e.g., the back end) of a gas turbine engine.
  • forward refers to the direction associated with the intake (e.g., the front end) of a gas turbine engine.
  • a first component that is "radially outward" of a second component means that the first component is positioned at a greater distance away from the engine central longitudinal axis than the second component.
  • a first component that is “radially inward” of a second component means that the first component is positioned closer to the engine central longitudinal axis than the second component.
  • a first component that is radially inward of a second component rotates through a circumferentially shorter path than the second component.
  • the terminology “radially outward” and “radially inward” may also be used relative to references other than the engine central longitudinal axis.
  • a first component that is "radially outward" of a second component means that the first component is positioned at a greater distance away from the engine central longitudinal axis than the second component.
  • distal refers to the direction outward, or generally, away from a reference component.
  • proximal refers to a direction inward, or generally, towards the reference component.
  • honeycomb structures are brazed onto the turbine vanes as part of the repair process during a maintenance shop visit (e.g., at an inner diameter (ID) or an outer diameter (OD) of a respective turbine vane.
  • braze tape is adhered onto the honeycomb structure prior to coupling the honeycomb structure to the turbine vane.
  • the honeycomb structure may then be tack welded onto the turbine vane prior to undergoing brazing in a furnace, in accordance with various embodiments.
  • methods and systems disclosed herein may provide significantly more consistent braze tape application and/or a controlled output of braze tape application to honeycomb structures compared to typical processes.
  • a braze tape adhered as disclosed herein may be applied evenly across the cells of the honeycomb structure, pulling in the braze tape all at once in a single iteration, in accordance with various embodiments.
  • the methods and systems may reduce ergonomic risks of typical manual processes that have repetitive manual force exerted manually by an individual performing a repair, in accordance with various embodiments.
  • the systems and methods disclosed herein may be scaled to allow operations of multiple honeycomb structures concurrently, vastly increasing process output. In various embodiments, overall productivity and safety may be achieved by the systems and methods disclosed herein.
  • Gas turbine engine 20 may be a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
  • fan section 22 can drive coolant (e.g., air) along a path of bypass airflow B while compressor section 24 can drive coolant along a core flow path C for compression and communication into combustor section 26 then expansion through turbine section 28.
  • coolant e.g., air
  • compressor section 24 can drive coolant along a core flow path C for compression and communication into combustor section 26 then expansion through turbine section 28.
  • Gas turbine engine 20 may generally comprise a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A-A' relative to an engine static structure 36 or engine case via several bearing systems 38, 38-1, and 38-2.
  • Engine central longitudinal axis A-A' is oriented in the Z direction on the provided X-Y-Z axes.
  • various bearing systems 38 at various locations may alternatively or additionally be provided, including for example, bearing system 38, bearing system 38-1, and bearing system 38-2.
  • Low speed spool 30 may generally comprise an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46.
  • Inner shaft 40 may be connected to fan 42 through a geared architecture 48 that can drive fan 42 at a lower speed than low speed spool 30.
  • Geared architecture 48 may comprise a gear assembly 60 enclosed within a gear housing 62.
  • Gear assembly 60 couples inner shaft 40 to a rotating fan structure.
  • High speed spool 32 may comprise an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54.
  • a combustor 56 may be located between high pressure compressor 52 and high pressure turbine 54.
  • a mid-turbine frame 57 of engine static structure 36 may be located generally between high pressure turbine 54 and low pressure turbine 46.
  • Mid-turbine frame 57 may support one or more bearing systems 38 in turbine section 28.
  • Inner shaft 40 and outer shaft 50 may be concentric and rotate via bearing systems 38 about the engine central longitudinal axis A-A', which is collinear with their longitudinal axes.
  • A-A' the engine central longitudinal axis A-A'
  • the core airflow may be compressed by low pressure compressor 44 then high pressure compressor 52, mixed and burned with fuel in combustor 56, then expanded over high pressure turbine 54 and low pressure turbine 46.
  • Turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
  • Gas turbine engine 20 may be, for example, a high-bypass ratio geared aircraft engine. In various embodiments, the bypass ratio of gas turbine engine 20 may be greater than about six (6). In various embodiments, the bypass ratio of gas turbine engine 20 may be greater than ten (10).
  • geared architecture 48 may be an epicyclic gear train, such as a star gear system (sun gear in meshing engagement with a plurality of star gears supported by a carrier and in meshing engagement with a ring gear) or other gear system. Geared architecture 48 may have a gear reduction ratio of greater than about 2.3 and low pressure turbine 46 may have a pressure ratio that is greater than about five (5). In various embodiments, the bypass ratio of gas turbine engine 20 is greater than about ten (10:1).
  • the diameter of fan 42 may be significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 may have a pressure ratio that is greater than about five (5:1).
  • Low pressure turbine 46 pressure ratio may be measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of low pressure turbine 46 prior to an exhaust nozzle.
  • a gas turbine engine 20 may comprise an industrial gas turbine (IGT) or a geared aircraft engine, such as a geared turbofan, or non-geared aircraft engine, such as a turbofan, or may comprise any gas turbine engine as desired.
  • IGT industrial gas turbine
  • a geared aircraft engine such as a geared turbofan
  • non-geared aircraft engine such as a turbofan
  • each of low pressure compressor 44, high pressure compressor 52, low pressure turbine 46, and high pressure turbine 54 in gas turbine engine 20 may comprise one or more stages or sets of rotating blades 101 and one or more stages or sets of stationary vanes 102 axially interspersed with the associated blade stages but non-rotating about engine central longitudinal axis A-A'.
  • the compressor and turbine sections 24, 28 may be referred to as rotor assemblies 110.
  • Each compressor stage and turbine stage may comprise multiple interspersed stages of blades 101 and vanes 102.
  • Within the rotor assemblies 110 of gas turbine engine 20 are multiple rotor disks, which may include one or more cover plates or minidisks.
  • FIG. 2 schematically shows, by example, a portion of an engine section 80, which is illustrated as a turbine section 28 of gas turbine engine 20. It will be understood that the repair systems and methods in the present disclosure are not limited to the turbine section 28 and could extend to other sections of the gas turbine engine 20, including but not limited to compressor section 24.
  • Engine section 80 may include alternating rows of blades 101 and vanes 102 comprising airfoils 100 that extend into the core flow path C.
  • the rotor assemblies 110 can carry a plurality of rotating blades 101, while each vane assembly can carry a plurality of vanes 102 that extend into the core flow path C.
  • Blades 101 create or extract energy (in the form of pressure) from the core airflow that is communicated through the gas turbine engine 20 along the core flow path C.
  • Vanes 102 direct the core airflow to the blades 101 to either add or extract energy.
  • Vanes 102 may be arranged circumferentially about engine central longitudinal axis A-A'.
  • a set of blades 101 may be coupled about a circumference of a generally circular disk 104, which may be disposed radially inward of core flow path C.
  • Disk 104 with blades 101 may comprise a rotor assembly 110 configured to rotate about engine central longitudinal axis A-A'.
  • Blades 101 and vanes 102 may generally be referred to as airfoils 100.
  • Each airfoil 100 illustrated as vane 102, has an airfoil body 120 having a leading edge 124 facing a forward direction in the gas turbine engine 20 and a trailing edge 126 facing an aft direction.
  • An airfoil 100 may include a pressure side wall (i.e. having a generally concave surface) and a suction side wall (i.e. having a generally convex surface) joined together at the respective leading edge 124 and trailing edge 126.
  • Each vane 102 may include an inner diameter (ID) platform 130 at an inner diameter end of the airfoil body 120 and an outer diameter (OD) platform 140 disposed at an OD end of the airfoil body 120.
  • the airfoil body 120 may extend radially outward from ID platform 130 at the inner diameter end of the airfoil body 120 to the OD platform 140 at the outer diameter end of the airfoil body 120.
  • Airfoil body 120, ID platform 130 and OD platform 140 may be integrally formed.
  • the term “integrated” or “integral” may include forming one, single continuous piece (i.e., a monolithic component). Casting may be used to form airfoils 100 of FIG. 2 .
  • the repair systems and methods, as described herein may be useful for any system or component that couples a braze component to a honeycomb structure prior to brazing, as described further herein.
  • the term "airfoil,” as used herein may refer to either a turbine blade 101 or a turbine vane 102, in accordance with various embodiments.
  • the repair system and method, as described herein may be used for other components of a gas turbine engine 20 with honeycomb structures.
  • engine section 80 further comprises a sealing system 150 between blades 101 and vanes 102.
  • the sealing system 150 may comprise a knife edge 152 extending radially outward from a flange 151 of blade 101 and a honeycomb seal land 154 disposed radially outward from the knife edge 152.
  • the knife edge 152 may be spaced apart radially from the honeycomb seal land 154 prior to operation and configured to contact the knife edge 152 during normal operation.
  • the honeycomb seal land 154 may act as a sacrificial material which the knife edge 152 abrades to maintain a desired tip clearance during operation of the gas turbine engine 20.
  • the repair systems and methods disclosed herein may be utilized to repair honeycomb seal land 154.
  • the honeycomb seal land 154 may be coupled to a radially inner surface of the ID platform 130 of vanes 102.
  • the honeycomb seal land 154 may be an annular structure extending circumferentially around the ID platform 130 of blades 101.
  • honeycomb structure 300 for use in a gas turbine engine (e.g., gas turbine engine 20 from FIG. 1 ), in accordance with various embodiments.
  • the honeycomb structure 300 may be coupled to an inner diameter surface of an ID platform of a vane assembly (e.g., ID platform 120 of vanes 102 from FIG. 2 ).
  • honeycomb structure 300 may be used in cryogenic rocket engine turbopumps as a damper seal at a balance piston location and in application in high-pressure centrifugal compressors, in accordance with various embodiments.
  • the honeycomb structure 300 may comprise a range of materials, such as stainless steel, aluminum, plastic, or the like.
  • the honeycomb structure 300 comprises a plurality of hexagonal cells 310. Each cell in the plurality of hexagonal cells 310 is disposed adjacent to an adjacent hexagonal cell in the plurality of hexagonal cells 310.
  • the honeycomb structure 300 may provide a material with minimal density (e.g., lower weight) and relatively high out-of-plane compression properties and out-out-of-plane shear properties.
  • honeycomb assembly 400 with a honeycomb structure 300 and a braze component 350 coupled thereto along section line A-A from FIG. 3A and manufactured in accordance with the repair systems and methods described further herein is illustrated, in accordance with various embodiments.
  • the honeycomb assembly 400 comprises the honeycomb structure 300 and the braze component 350.
  • Each cell in the plurality of hexagonal cells 310 extends from a first end 312 to a second end 314.
  • Each cell in the plurality of hexagonal cells 310 defines a hexagonal cavity therein.
  • the braze component 350 is adhered to the honeycomb structure 300 via system 500 illustrated in FIGs. 4A and 4B and manufactured in accordance with method 600 from FIG. 5 .
  • the braze component 350 comprises braze tape.
  • the braze component 350 may comprise a polymer bonded flexible tape comprising a nickel-based super alloy or a cobalt based super alloy and a brazing filler metal power.
  • the braze component 350 is flexible and configured for a diffusion brazing process.
  • the honeycomb assembly 400 made in accordance with system 500 and method 600 as described further herein provides a more consistent and controlled honeycomb assembly 400 relative to typical manual systems and processes for applying a braze component 350 to a honeycomb assembly 400.
  • the braze component 350 may be disposed more uniformly and at a more consistent depth relative to the second end 314 compared to a typical manually applied braze component 350.
  • "more uniformly” refers to a standard deviation in local density per square inch being smaller relative to a typical standard deviation in local density per square inch of the braze component 350.
  • braze component 350 refers to a standard deviation in local depth of the braze component 350 measured periodically being smaller relative to a typical standard deviation of typical honeycomb assemblies applied via manual processes. Although illustrated as being a depth below the second end 314, the present disclosure is not limited in this regard.
  • the braze component 350 may be substantially flush with the second end 314, or extend above the second end 314, in accordance with various embodiments.
  • FIGS. 4A and 4B a cross-sectional view of a system 500 for applying a braze component 350 to a honeycomb structure 300 to form a honeycomb assembly 400 in accordance with FIG. 3B is illustrated, in accordance with various embodiments.
  • the system 500 comprises a vacuum device 510, a porous platform 520, and a cover 530.
  • the porous platform 520 is disposed within the vacuum device 510.
  • the porous platform 520 is configured to receive a honeycomb structure 300 thereon.
  • a vacuum device 510 may be separate from a housing that houses the vacuum device therein and is configured to create an air-tight chamber for the honeycomb structure 300 and braze component 350 to be disposed within.
  • a braze component 350 may be placed on a second end 314 of the honeycomb structure 300 prior to placing the cover 530 thereon.
  • the cover 530 is placed over the honeycomb structure 300 to create an air-tight chamber and/or allow a suction from the vacuum device 510 to function (i.e., for the suction from the vacuum device 510 to pull the cover 530 towards the vacuum device 510, and also creating a force on the braze component 350 directed towards the honeycomb structure 300).
  • the porous platform 520 and the honeycomb structure 300 allows the vacuum device 510 to be in fluid communication with an internal surface of the braze component 350. In this regard, suction from the vacuum device 510 pulls the internal surface of the braze component 350 towards the suction as illustrated in FIG. 4B .
  • the vacuum device 510 may be any vacuum pump known in the art, such as a positive displacement pump, a momentum transfer pump, a regenerative pump, or the like.
  • the cover 530 and the vacuum device 510 may be sealed when performing method 600 described further herein.
  • the cover 530 may be pulled towards a first end 512 of the vacuum device 510 in response to the vacuum device 510 producing a suction towards the first end 512 as shown in FIG. 4B .
  • the honeycomb structure 300 may be arcuate in shape and configured for a portion of a honeycomb seal land 154 from FIG. 2 .
  • the honeycomb structure 300 may be a flat component. In this regard, a shape of the honeycomb structure 300 is not limited by the present disclosure.
  • the system 500 may further include a weight placed on top of cover 530 to further apply pressure of the braze component 350 into the plurality of hexagonal cells 310 of the honeycomb structure 300.
  • a balance between suction force and/or an external weight may provide a consistent and controlled application of braze component 350 to the honeycomb structure 300, in accordance with various embodiments.
  • the cover 530 may be configured to be heated during application of the braze component 350 to the honeycomb structure 300.
  • heating the braze component 350 may soften the braze component 350 making it easier to adhere the braze component 350 to the honeycomb structure 300.
  • the heating device may include an electrical resistance heater.
  • the electrical resistance heater may be integral with the cover 530, spaced apart from the cover 530, or the like.
  • the system 500 may further comprise a controller 540.
  • the controller 540 may be in operable communication with the vacuum device 510.
  • the controller 540 may be operable to command the vacuum device 510 to supply at least a partial vacuum, in accordance with various embodiments.
  • the method 600 comprises disposing a honeycomb structure on a porous platform within a vacuum device (step 602).
  • the honeycomb structure may be in accordance with honeycomb structure 300 from FIGs. 3-4B .
  • the method 600 further comprises disposing a braze component on a second end of the honeycomb structure (step 604).
  • the second end is opposite a first end of the honeycomb structure, defining a plurality of hexagonal cells therebetween.
  • the method 600 further comprises creating an air-tight chamber between a cover and the vacuum device, the honeycomb structure and the braze component being disposed within the air-tight chamber (step 606).
  • the method 600 further comprises pulling the cover towards the suction device in response to the suction device creating at least a partial vacuum within the air-tight chamber (step 608).
  • a suction force created by the partial vacuum may be supplied relatively evenly across a plurality of cells of the honeycomb structure.
  • the braze component may be more consistent and controllable relative to typical systems, in accordance with various embodiments.
  • the suction force may be calibrated to create an optimal pulling force of the braze component into a plurality of cells of the honeycomb structure.
  • the method 600 further comprises adhering the braze component to the honeycomb structure within a plurality of hexagonal cells of the honeycomb structure in response to the cover applying a force against the braze component (step 610).
  • the cover may be heated during steps 608 and 610.
  • an electrical resistance heater may supply a current to the cover, in accordance with various embodiments.
  • the braze component may be softer and adhere more easily than without heating the cover.
  • a weight may be placed on the cover prior to applying the partial vacuum in step 608.
  • method 600 may produce the honeycomb assembly 400 from FIG. 3B .
  • a method of repairing a honeycomb seal land may comprise the method 600, tack welding a resultant honeycomb assembly (e.g., honeycomb assembly 400 to an inner diameter surface of a platform (e.g., ID platform 130 from FIG. 2 ), and brazing the platform and the resultant honeycomb assembly.
  • tack welding may be performed on the second end 314 of the honeycomb structure 300 proximate the braze component 350.
  • an open face honeycomb structure may be repaired and configured to be disposed opposite a knife edge (e.g., knife edge 152 from FIG. 2 ), in accordance with various embodiments.
  • references to "various embodiments”, “one embodiment”, “an embodiment”, “an example embodiment”, etc. indicate that the embodiment described may include a particular feature, structure, or characteristic, but every embodiment may not necessarily include the particular feature, structure, or characteristic. Moreover, such phrases are not necessarily referring to the same embodiment. Further, when a particular feature, structure, or characteristic is described in connection with an embodiment, it is submitted that it is within the knowledge of one skilled in the art to affect such feature, structure, or characteristic in connection with other embodiments whether or not explicitly described. After reading the description, it will be apparent to one skilled in the relevant art(s) how to implement the disclosure in alternative embodiments.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Sealing Using Fluids, Sealing Without Contact, And Removal Of Oil (AREA)
EP22155874.5A 2021-02-26 2022-02-09 Procédés de réparation et systèmes de structures en nid d'abeille dans des moteurs à turbine à gaz Pending EP4063062A1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US17/187,008 US11370070B1 (en) 2021-02-26 2021-02-26 Repair methods and systems for honeycomb structures in gas turbine engines

Publications (1)

Publication Number Publication Date
EP4063062A1 true EP4063062A1 (fr) 2022-09-28

Family

ID=80682221

Family Applications (1)

Application Number Title Priority Date Filing Date
EP22155874.5A Pending EP4063062A1 (fr) 2021-02-26 2022-02-09 Procédés de réparation et systèmes de structures en nid d'abeille dans des moteurs à turbine à gaz

Country Status (2)

Country Link
US (2) US11370070B1 (fr)
EP (1) EP4063062A1 (fr)

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3001274A (en) * 1957-01-18 1961-09-26 Solar Aircraft Co Brazing article and method
US3844027A (en) * 1973-06-25 1974-10-29 Chrysler Corp Copper brazing of matrix structures
EP2612951A2 (fr) * 2012-01-05 2013-07-10 General Electric Company Procédé de fabrication d'un joint cellulaire
EP3299114A1 (fr) * 2016-09-21 2018-03-28 General Electric Company Gel de brasage, procédé de brasage et article brasé
KR102097612B1 (ko) * 2019-04-16 2020-04-06 하나아이티엠(주) 저압 터빈부 에어씰 브레이징 파우더 도포 장치를 이용한 브레이징 공법
US20200316701A1 (en) * 2017-11-08 2020-10-08 Safran Nacelles Method for manufacturing a structural and/or acoustic panel for a nacelle of an aircraft propulsion unit, and corresponding device

Family Cites Families (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2962811A (en) * 1955-09-19 1960-12-06 Rohr Aircraft Corp Method of making stainless steel honeycomb panels
US3046648A (en) * 1959-04-13 1962-07-31 Aircraft Prec Products Inc Method of manufacturing replaceable labyrinth type seal assembly
US3173813A (en) * 1960-03-28 1965-03-16 Douglas Aircraft Co Inc Method and apparatus for bonding stainless steel honeycomb
US3091684A (en) * 1961-07-13 1963-05-28 Gen Dynamics Corp Brazing apparatus
US3412917A (en) * 1965-10-21 1968-11-26 Northrop Corp Bonding or brazing apparatus
US3656224A (en) * 1969-04-09 1972-04-18 Rohr Corp Method of forming a honeycomb core panel
US3737978A (en) * 1970-01-07 1973-06-12 Aeronca Inc Brazing method
US3722071A (en) 1971-09-30 1973-03-27 Aeronca Inc Brazing powder deposition method
US4449714A (en) 1983-03-22 1984-05-22 Gulf & Western Industries, Inc. Turbine engine seal and method for repair thereof
US5702050A (en) * 1995-04-28 1997-12-30 Mitsubishi Jukogyo Kabushiki Kaisha Method of brazing a honeycomb
US20060261136A1 (en) * 2005-05-17 2006-11-23 Calsonic Kansei Corporation Diffusion bonding method for forming metal substrate

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3001274A (en) * 1957-01-18 1961-09-26 Solar Aircraft Co Brazing article and method
US3844027A (en) * 1973-06-25 1974-10-29 Chrysler Corp Copper brazing of matrix structures
EP2612951A2 (fr) * 2012-01-05 2013-07-10 General Electric Company Procédé de fabrication d'un joint cellulaire
EP3299114A1 (fr) * 2016-09-21 2018-03-28 General Electric Company Gel de brasage, procédé de brasage et article brasé
US20200316701A1 (en) * 2017-11-08 2020-10-08 Safran Nacelles Method for manufacturing a structural and/or acoustic panel for a nacelle of an aircraft propulsion unit, and corresponding device
KR102097612B1 (ko) * 2019-04-16 2020-04-06 하나아이티엠(주) 저압 터빈부 에어씰 브레이징 파우더 도포 장치를 이용한 브레이징 공법

Also Published As

Publication number Publication date
US20230061452A1 (en) 2023-03-02
US11819966B2 (en) 2023-11-21
US11370070B1 (en) 2022-06-28

Similar Documents

Publication Publication Date Title
US11519293B2 (en) Cleaning system and a method of cleaning
WO2013180954A1 (fr) Système d'alimentation en réfrigérant pour turbine à haute pression
EP3252271B1 (fr) Ensemble pour aube, étage de soufflante et moteur à turbine à gaz associés
US10934845B2 (en) Dual cooling airflow to blades
EP3315732A1 (fr) Dosage d'air de refroidissement pour joints d'étanchéité de fente extérieurs d'aube
EP3572214A1 (fr) Outil et procédés de liaison de surface portante
US11370070B1 (en) Repair methods and systems for honeycomb structures in gas turbine engines
US20190323359A1 (en) Blade and a method of manufacturing a blade
US10513947B2 (en) Adjustable flow split platform cooling for gas turbine engine
EP3299587B1 (fr) Aube de moteur à turbine à gaz
US10385717B2 (en) Multi-ply seal
US9926789B2 (en) Flow splitting baffle
EP3415716B1 (fr) Refroidissement des extrémités des aubes de turbine
EP3693542B1 (fr) Arbre métallique
US11092033B2 (en) Method and tooling for manufacturing a fan case assembly
EP4357587A1 (fr) Rotor à aubage intégral pour un moteur à turbine à gaz, moteur à turbine à gaz, et procédé de fabrication d'un rotor à aubage intégral d'un moteur à turbine à gaz
US20210270140A1 (en) An Aerofoil Structure and a Method of Manufacturing an Aerofoil Structure for a Gas Turbine Engine
US10415416B2 (en) Fluid flow assembly

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: THE APPLICATION HAS BEEN PUBLISHED

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: REQUEST FOR EXAMINATION WAS MADE

17P Request for examination filed

Effective date: 20230328

RBV Designated contracting states (corrected)

Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR

RAP3 Party data changed (applicant data changed or rights of an application transferred)

Owner name: RTX CORPORATION